Composite barrel sections for aircraft fuselages and other structures

ABSTRACT

Composite sections for aircraft fuselages and methods and systems for manufacturing such sections are herein. A composite aircraft panel manufacturing assembly comprises a skin, a stiffener, a tubular bladder, and at least one ply of fabric. The stiffener comprises first and second flange portions bonded to the skin. The stiffener further comprises an interior surface, spaced apart from the skin, and extending between the first and second flange portions. The tubular bladder extends longitudinally between the interior surface of the stiffener and the skin. The at least one ply of fabric extends around the bladder and is bonded to the interior surface of the stiffener and an adjacent portion of the skin between the first and second flange portions.

CROSS-REFERENCE TO RELATED APPLICATION

The present application is a divisional application of U.S. patentapplication Ser. No. 12/236,355, filed Sep. 23, 2008 now U.S. Pat. No.8,157,212, status allowed, which is a continuation of U.S. applicationSer. No. 10/851,381 now U.S. Pat. No. 7,527,222, filed May 20, 2004,which claims priority to U.S. Provisional Patent Application No.60/559,890 filed Apr. 6, 2004, and incorporated herein in its entiretyby reference; and this application is related to U.S. patent applicationSer. No. 12/020,956, filed Jan. 28, 2008, which is a divisionalapplication of U.S. Pat. No. 7,527,222, filed May 20, 2004, andincorporated herein in its entirety by reference.

TECHNICAL FIELD

The following disclosure relates generally to aircraft structures and,more particularly, to composite barrel sections for aircraft fuselagesand methods and systems for manufacturing such barrel sections.

BACKGROUND

Aircraft manufacturers continually strive for ways to increase aircraftperformance and reduce manufacturing costs. One well-known method forincreasing aircraft performance is to reduce airframe weight through theuse of composite materials having relatively high strength-to-weightratios. Composite materials have been used on airframes for fighteraircraft, high-performance private aircraft, and business jets. Largeraircraft, however, such as large commercial transport aircraft,typically use metallic materials for all or most of the primarystructure. The fuselage shells for commercial transport aircraft, forexample, are typically manufactured from aluminum and other metals.

Conventional methods for manufacturing business jet airframes withcomposite materials typically require extensive tooling fixtures andlabor-intensive assembly procedures. One known method used by theRaytheon Aircraft Company of Wichita, Kans., to manufacture the PremierI and Hawker Horizon business jets involves wrapping carbon fibersaround a rotating mandrel with an automated fiber placement system. Themandrel provides the basic shape of a fuselage section. The carbonfibers are preimpregnated with a thermoset epoxy resin, and they areapplied over the rotating mandrel in multiple plies to form an interiorskin of the fuselage section. The interior skin is then covered with alayer of honeycomb core. The fiber placement system then appliesadditional plies of preimpregnated carbon fibers over the honeycomb coreto form an exterior skin that results in a sandwich structure. The finalply includes a hybrid fabric of carbon fiber and fine metallic wires toprovide lightning strike protection.

The Premier I fuselage includes two composite fuselage sections formedin the foregoing manner. The Hawker Horizon fuselage includes threesections formed in this manner. After forming, the respective fuselagesections are bonded together along circumferential joints to form thecomplete fuselage shell. Another method for forming composite fuselageshells in accordance with the prior art involves forming fuselage halvesor quarter-panels separately (for example, by a fabric lay-up process),and then joining the separate parts together along longitudinal jointsto form a complete fuselage cross-section.

Filament winding, fiber placement, and tape laying are three knownmethods for applying unidirectional composite fibers to a rotatingmandrel to form a continuous cylindrical skin. In a filament windingprocess, the mandrel is typically suspended horizontally between endsupports. The mandrel rotates about the horizontal axis as a fiberapplication instrument moves back and forth along the length of themandrel, placing fiber onto the mandrel in a predeterminedconfiguration. In most applications, the filament winding apparatuspasses the fiber material through a resin “bath” just before thematerial touches the mandrel. This is called “wet winding.” In otherapplications, the fiber has been preimpregnated with resin, eliminatingthe need for the resin bath. Following oven or autoclave curing of theresin, the mandrel can remain in place and become part of the woundcomponent, or it can be removed.

The fiber placement process typically involves the automated placementof multiple “tows” (i.e., untwisted bundles of continuous filaments,such as carbon or graphite fibers, preimpregnated with a thermoset resinmaterial such as epoxy) tape, or slit tape onto a rotating mandrel athigh speed. A typical tow is between about 0.12″ and 0.25″ wide whenflattened. Conventional fiber placement machines dispense multiple towsto a movable payoff head that collimates the tows (i.e., renders thetows parallel) and applies the tows to the rotating mandrel surfaceusing one or more compaction rollers that compress the tows against thesurface. In addition, such machines typically include means fordispensing, clamping, cutting and restarting individual tows duringplacement.

Tape laying is similar to the fiber placement process described aboveexcept that preimpregnated fiber tape, rather than individual tows, islaid down on the rotating mandrel to form the part. One form of tapeincludes a paper backing that maintains the width and orientation of thefibers. The paper backing is removed during application. Slit tape istape that has been slit after being produced in standard widths by themanufacturer. Slitting the tape results in narrower widths that allowenhanced steerability and tailoring during application to achieveproducibility and design objectives. Slit tape can have widths varyingfrom about 0.12 inch up to about 6 inches, and may or may not includebacking paper. Another form of tape includes multiple individual fiberswoven together with a cloth material. As used throughout thisdisclosure, unless otherwise indicated, the term “tape” refers to tape,tape with backing paper, slit tape, and other types of compositematerial in tape form for use in manufacturing composite structures.Tape laying is often used for parts with highly complex contours orangles because the tape allows relatively easy directional changes.

SUMMARY

The present invention is directed generally toward composite sectionsfor aircraft fuselages and other structures. A section configured inaccordance with one aspect of the invention includes a skin having aplurality of fiber tows forming a continuous surface extending 360degrees about an axis. The section can further include at least firstand second stiffeners. The first stiffener can have a first flangeportion bonded to an interior surface of the skin and a first raisedportion projecting inwardly and away from the interior surface of theskin. The second stiffener can have a second flange portion bonded tothe interior surface of the skin and a second raised portion projectinginwardly and away from the interior surface of the skin. A sectionconfigured in accordance with another aspect of the invention caninclude a skin having a plurality of fiber tapes forming the continuoussurface instead of or in addition to the plurality of collimated fibertows.

A method for manufacturing a section of a fuselage in accordance withone aspect of the invention includes positioning a plurality ofstiffeners on a mandrel assembly and rotating the mandrel assembly abouta longitudinal axis. The method can further include applying a pluralityof fiber tows to form a continuous skin extending 360 degrees around themandrel assembly. After application of the fiber tows, the stiffenersand the fiber tows can be cocured. A method for manufacturing a sectionof a fuselage in accordance with another aspect of the invention caninclude laying fiber tape over the stiffeners on the rotating mandrelassembly instead of or in addition to the fiber tows.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a partially hidden isometric view of an aircraft having afuselage that includes a plurality of barrel sections configured inaccordance with an embodiment of the invention.

FIGS. 2A and 2B are an exploded isometric view and an assembledisometric view, respectively, of a portion of a fuselage barrel sectionconfigured in accordance with an embodiment of the invention.

FIGS. 3A and 3B are top and end views, respectively, of a portion of afuselage barrel section configured in accordance with another embodimentof the invention.

FIGS. 4A and 4B are top and end views, respectively, of a portion of afuselage barrel section configured in accordance with a furtherembodiment of the invention.

FIGS. 5A and 5B are cross-sectional end views of portions of fuselagebarrel sections configured in accordance with yet other embodiments ofthe invention.

FIG. 6 is a partially schematic isometric view of a barrel sectionmanufacturing system configured in accordance with an embodiment of theinvention.

FIGS. 7A and 7B are enlarged, partially schematic isometric views of abarrel section stiffener loading station illustrating two stages of amethod for loading stiffeners onto a tool assembly in accordance with anembodiment of the invention.

FIG. 8 is an enlarged, partially schematic isometric view of a barrelsection laminating station configured in accordance with an embodimentof the invention.

FIG. 9 is an enlarged, partially schematic isometric view of a barrelsection vacuum-bagging station configured in accordance with anembodiment of the invention.

FIG. 10 is an enlarged, partially schematic isometric view of a barrelsection curing station configured in accordance with an embodiment ofthe invention.

FIG. 11 is an enlarged, partially schematic isometric view of a barrelsection inspection station configured in accordance with an embodimentof the invention.

FIG. 12 is an enlarged, partially schematic isometric view of a barrelsection trimming station configured in accordance with an embodiment ofthe invention.

FIG. 13 is an enlarged, partially schematic isometric view of a barrelsection assembly station configured in accordance with an embodiment ofthe invention.

FIGS. 14A-14C are cross-sectional end views illustrating stages of amethod for bonding a stiffener to a laminate in accordance with anembodiment of the invention.

DETAILED DESCRIPTION

The following disclosure describes composite barrel sections foraircraft fuselages and other structures, and methods and systems formanufacturing such barrel sections. Throughout this disclosure, the termbarrel section is used for convenience to refer generally to an enclosedshell structure extending 360 degrees about an axis. Such structures caninclude, for example, cylindrical shells having circular, oval,elliptical, egg-shaped, and other symmetrical and/or asymmetricalcross-sectional shapes. Such structures can further include enclosed,non-cylindrical shells. Certain details are set forth in the followingdescription and in FIGS. 1-14C to provide a thorough understanding ofvarious embodiments of the invention. Other details describingwell-known structures and systems often associated with aircraftstructures and composite fabrication techniques are not set forth in thefollowing disclosure to avoid unnecessarily obscuring the description ofthe various embodiments of the invention.

Many of the details, dimensions, angles, and other features shown in theFigures are merely illustrative of particular embodiments of theinvention. Accordingly, other embodiments can have other details,dimensions, angles, and features without departing from the spirit orscope of the present invention. In addition, further embodiments can bepracticed without several of the details described below.

In the Figures, identical reference numbers identify identical or atleast generally similar elements. To facilitate the discussion of anyparticular element, the most significant digit or digits of anyreference number refer to the Figure in which that element is firstintroduced. For example, element 110 is first introduced and discussedwith reference to FIG. 1.

FIG. 1 is a partially hidden isometric view of an aircraft 100 having afuselage 102 that includes a plurality of barrel sections 110 configuredin accordance with an embodiment of the invention. In one aspect of thisembodiment described in greater detail below, each of the barrelsections 110 can be individually fabricated as a one-piece section fromcomposite materials, such as carbon fiber and/or graphite-epoxymaterials. After fabrication, the barrel sections 110 can be joinedtogether by adhesive bonding and/or mechanical fastening alongcircumferential joints 112 to form the fuselage 102.

In another aspect of this embodiment, the fuselage 102 can include apassenger cabin 104 configured to hold a plurality of passenger seats106. In the illustrated embodiment, the passenger cabin 104 isconfigured to hold at least about 50 of the passenger seats 106, e.g.,from about 50 to about 700 passenger seats. In another embodiment, thepassenger cabin 104 can be configured to hold from about 150 to about400 of the passenger seats 106. In other embodiments, the passengercabin 104 can be configured to hold more or fewer seats or,alternatively, the passenger seats 106 can be omitted and the cabinspace can be used for other purposes, such as hauling cargo.

FIG. 2A is an enlarged, partially exploded, interior isometric view of aportion of one of the barrel sections 110 of FIG. 1, configured inaccordance with an embodiment of the invention. FIG. 2B is an assembledisometric view of the barrel section portion of FIG. 2A. Referring toFIGS. 2A and 2B together, the barrel section 110 can include a pluralityof stiffeners 230 (identified individually as stiffeners 230 a-d)attached to a skin 220. Each of the stiffeners 230 can include a raisedportion 234 projecting away from the skin 220 and a plurality of flangeportions 231 (identified as a plurality of first flange portions 231 aextending outwardly from one side of the stiffener 230, and a pluralityof second flange portions 231 b extending outwardly from an oppositeside of the stiffener 230). The flange portions 231 can be mateddirectly to the skin 220. In the illustrated embodiment, the stiffeners230 have hat-shaped cross-sections. In other embodiments describedbelow, however, the stiffeners 230 can have other cross-sectionalshapes.

In one embodiment described in greater detail below, the skin 220 andthe stiffeners 230 can include composite materials, such as carbon fibermaterials. In this embodiment, the stiffeners 230 can be bonded to theskin 220. For example, in one embodiment described in detail below, thestiffeners 230 can be bonded to the skin 220 during a cocuring processin which the stiffeners 230 and the skin 220 are cocured at an elevatedtemperature and pressure. In another embodiment, the stiffeners 230 canbe pre-cured and adhesively bonded to the skin 220 when exposed to anelevated temperature and pressure. In yet other embodiments, thestiffeners 230 can be mechanically fastened to the skin 220.

Each of the stiffeners 230 can be positioned on the skin 220 so that theplurality of first flange portions 231 a of one stiffener 230 arealigned with the corresponding plurality of second flange portions 231 bof an adjacent stiffener 230. For example, each of the first flangeportions 231 a can include a first outer edge 233 a, and each of thesecond flange portions 231 b can include a corresponding second outeredge 233 b. In one embodiment, the first outer edge 233 a can be spacedapart from the second outer edge 233 b by a distance D of about 0.5 inchor less. In another embodiment, the distance D can be about 0.2 inch orless, e.g., about 0.1 inch. In yet another embodiment, the stiffeners230 can be positioned on the skin 220 such that the first flangeportions 231 a at least approximately contact the second flange portions231 b. In this case, the distance D is at least approximately zero. Whenthe flange portions 231 are aligned in the foregoing manner, the flangeportions 231 can form a plurality of at least approximately continuoussupport surfaces 235 extending between the raised portions 234 of thestiffeners 230.

The barrel section 110 can further include a plurality of supportmembers or frames 240 (identified individually as a first frame 240 aand a second frame 240 b). In the illustrated embodiment, the frames 240are two-piece frames that include a first frame section 241 and a secondframe section 242. In this embodiment, the second frame section 242 hasa C-shaped cross-section. In other embodiments, the second frame section242 can have other cross-sectional shapes, such as an L-shapedcross-section. In yet other embodiments, the frames 240 can be omittedor, alternatively, the barrel section 110 can include other framescomposed of more or fewer frame sections.

The first frame section 241 includes a base portion 244 and anupstanding portion 246 projecting away from the base portion 244. Theupstanding portion 246 can include a plurality of openings, e.g., “mouseholes” 248 through which the raised portions 234 of the stiffeners 230extend. The base portion 244 can include a plurality of mating surfaces243 extending between the mouse holes 248. The mating surfaces 243 areconfigured to contact corresponding ones of the support surfaces 235extending between the raised portions 234 of the stiffeners 230. Themating surfaces 243 of the illustrated embodiment are absent any jogglesbetween the mouse holes 248 because the corresponding support surfaces235 to which they mate are at least approximately continuous between thestiffeners 230 and do not include any significant surface steps ormisalignments. An advantage of this feature is that it avoids the addedcosts associated with manufacturing frames with joggles. Such costs maybe particularly significant when working with composite materialsbecause, unlike creating joggles or steps in metals, which are malleableand can be easily formed, creating joggles or steps in compositesurfaces typically requires special tooling and/or post-cure machining.

In one embodiment of the invention, the first frame section 241 can beattached to the barrel section 110 first, and then the second framesection 242 can be attached to the first frame section 241. Whenattaching the first frame section 241 to the barrel section 110, thebase portion 244 of the first frame section 241 is mated to the flangeportions 231 of the stiffeners 230 without being mated to the skin 220.That is, the mating surfaces 243 of the base portion 244 contact thesupport surfaces 235 but not the skin 220. In this manner, the flangeportions 231 are effectively sandwiched between the first frame section241 and the skin 220. In one embodiment, the first frame section 241 canbe fastened to the barrel section 110 with a series of suitablefasteners 252, as shown in FIG. 2B. In another embodiment, the baseportion 244 can be adhesively bonded directly to the flange portions231.

After the first frame section 241 has been attached to the barrelsection 110, the second frame section 242 can be attached to the firstframe section 241. In one embodiment, the second frame section 242 canbe fastened to the upstanding portion 246 of the first frame section 241with a series of suitable fasteners 250, as shown in FIG. 2A. In anotherembodiment, the second frame section 242 can be adhesively bonded to theupstanding portion 246. One advantage of attaching the second framesection 242 to the first frame section 241 after the first frame section241 has been installed is that the final position of the second framesection 242 can be adjusted to compensate for any misalignment of thefirst frame section 241 that may have occurred during installation ofthe first frame section 242. In other embodiments, however, the firstframe section 241 can be attached to the second frame section 242 first,and then the frame 240 can be attached to the barrel section 110 as acomplete unit.

In another embodiment of the invention, the flange portions 231 of thestiffeners 230 can be at least partially omitted. In this embodiment, araised portion can be formed on the skin 220 between the stiffeners 230with an additional ply or plies of material. The raised portion can takethe place of the flange portions 231 in forming the support surface 235to which the base portion 244 of the first frame section 241 mates.

FIGS. 3A and 3B are top and end views, respectively, of a portion of abarrel section 310 configured in accordance with another embodiment ofthe invention. Referring to FIGS. 3A and 3B together, the barrel section310 can include a plurality of first stiffeners 336 and a plurality ofsecond stiffeners 338 attached to a skin 320. Each of the stiffeners 336and 338 can include a raised portion 334 projecting away from the skin320. Each of the first stiffeners 336 can further include a first flangeportion 337 a and an opposing second flange portion 337 b that are atleast generally straight. Each of the second stiffeners 338, however,can further include a plurality of first flange portions 331 a and aplurality of opposing second flange portions 33 b that extend outwardlyfrom the raised portion 334 to at least proximate corresponding flangeportions 337 of the adjacent first stiffeners 336. A frame (not shown)can mate to the flange portions 331 and 337 as described above withreference to FIGS. 2A and 2B.

FIGS. 4A and 4B are top and end views, respectively, of a portion of abarrel section 410 configured in accordance with a further embodiment ofthe invention. Referring to FIGS. 4A and 4B together, the barrel section410 can include a plurality of asymmetric stiffeners 450 attached to askin 420. Each of the asymmetric stiffeners 450 can include a pluralityof first flange portions 431 extending outwardly from one side of araised portion 434, and a second flange portion 437 extending outwardlyfrom an opposite side of the raised portion 434. The second flangeportion 437 can be at least approximately straight. The first flangeportions 431, however, can project outwardly from the raised portion 434to at least proximate the corresponding second flange portion 437 of theadjacent stiffener 450. A frame (not shown) can mate to the flangeportions 431 and 437 as described above with reference to FIGS. 2A and2B. FIGS. 5A and 5B are cross-sectional end views of portions of barrelsections 510 a and 510 b, respectively, configured in accordance withother embodiments of the invention. Referring first to FIG. 5A, in oneaspect of this embodiment, the barrel section 510 a includes a pluralityof I-section stiffeners 530 a attached to a skin 520 a. Each of theI-section stiffeners 530 a can include a plurality of first flangeportions 531 a and a plurality of second flange portions 531 b that areat least generally similar in structure and function to thecorresponding flange portions 231 described above with reference toFIGS. 2A and 2B. In another aspect of this embodiment, a frame 540 a canmate to the flange portions 531 as described above with reference toFIGS. 2A and 2B.

Referring next to FIG. 5B, in one aspect of this embodiment, the barrelsection 510 b includes a plurality of C-section stiffeners 530 battached to a skin 520 b. The C-section stiffeners 530 b can includeflange portions 531 that are at least generally similar in structure andfunction to the first flange portions 431 described above with referenceto FIGS. 4A and 4B. In another aspect of this embodiment, a frame 540 bcan mate to the flange portions 531 as described above with reference toFIGS. 2A and 2B.

FIG. 6 is a partially schematic isometric view of a barrel sectionmanufacturing system 600 arranged on a factory floor 602 in accordancewith an embodiment of the invention. In one aspect of this embodimentdescribed in greater detail below, the barrel section manufacturingsystem 600 includes a serial arrangement of manufacturing stationsconfigured to manufacture the fuselage barrel sections described abovewith reference to FIGS. 1-5B. As an overview, in the illustratedembodiment, barrel section fabrication begins at a stiffener loadingstation 610 before moving to a skin laminating station 620. After skinlamination, the barrel section (not shown) moves to a vacuum station 630for vacuum-bagging before moving to a curing station 640. From there,the barrel section moves successively to an inspection station 650, atrimming station 660, and an assembly station 670.

The foregoing arrangement of manufacturing stations is but onearrangement that can be used to manufacture the fuselage barrel sectionsdescribed above. In other embodiments, other manufacturing arrangementsand/or other types of manufacturing stations can be used in place of orin addition to one or more of the manufacturing stations illustrated inFIG. 6. For example, in one embodiment, one or more of the manufacturingstations can be positioned in a parallel arrangement rather than theserial-type arrangement illustrated in

FIG. 6. In another embodiment, two or more of the manufacturing stationscan be combined to form a single station.

FIGS. 7A and 7B are enlarged, partially schematic isometric views of thestiffener loading station 610 illustrating two stages of a method forloading a plurality of stiffeners 730 onto a barrel section toolassembly 700 in accordance with an embodiment of the invention.Referring first to FIG. 7A, in one aspect of this embodiment, the barrelsection tool assembly 700 includes a rotatable tool fixture 702configured to support a plurality of tool segments 706 (identifiedindividually as tool segments 706 a-f) in a cylindrical arrangement. Thetool segments 706 can be manufactured from a plurality of suitablematerials including steel, invar, aluminum, or composites. Each of thetool segments 706 can include a plurality of stiffener grooves 708configured to individually receive a corresponding one of the stiffeners730. In one embodiment, the stiffeners 730 can be hat-section stiffeners(e.g., hat section stiffeners that are at least generally similar instructure and function to the stiffeners 230 described above withreference to FIGS. 2A and 2B). In this embodiment, each of thestiffeners 730 is inverted in the corresponding stiffener groove 708 sothat the stiffener flange portions (e.g., the flange portions 231 ofFIG. 2A) lie in corresponding recesses formed in the tool segment 706adjacent to the stiffener grooves 708.

In another aspect of this embodiment, the stiffeners 730 can be leastgenerally uncured when placed in the stiffener grooves 708. In theuncured condition, the stiffeners 730 are relatively flimsy. As aresult, suitable tooling (not shown) may be required to at leasttemporarily hold the stiffeners 730 in position against the toolsegments 706 after installation in the stiffener grooves 708. In otherembodiments, the stiffeners 730 can be at least partially cured, inwhich case less or different tooling may be required to hold thestiffeners 730 in position.

Once the tool segments 706 are fully loaded with the stiffeners 730, thetool segments 706 are loaded onto the tool fixture 702, as illustratedin FIG. 7B. In one aspect of this embodiment, the tool fixture 702 isrotatably supported in a tool support structure 704 by a plurality ofrollers 705. The rollers 705 enable the tool fixture 702 to rotate abouta longitudinal axis 707. To prevent the stiffeners 730 from falling outof the stiffener grooves 708 during rotation, an innermost ply 721 ofcomposite fabric can be wrapped around the tool segments 706 to hold thestiffeners 730 in position. In other embodiments, the innermost ply 721can be omitted and the stiffeners 730 can be held in position by othermeans, including local tooling clips or other features. After theinnermost ply 721 has been fully installed, the tool support structure704 transports the tool assembly 700 to the laminating station 620 (FIG.6) via floor tracks 712.

The tool assembly 700 described above with reference to FIGS. 7A and 7Bis but one type of tool assembly that can be used in accordance with thepresent invention to position stiffeners in a cylindrical arrangementprior to the application of composite skin materials. In otherembodiments, other types of tool assemblies can be used. For example, inanother embodiment, a similar tool assembly can utilize a centralspindle for supporting and rotating the tool fixture 702 in place of theexternal rollers 705. In a further embodiment, the individual toolsegments 706 can be omitted and instead the tool fixture 702 can includea complete cylindrical surface configured to hold the stiffeners 730.This particular approach may offer the advantage of reduced stiffenerloading time. However, the other approach of using multiple toolsegments may have the advantage of reducing the time required toseparate the finished barrel section from the tool assembly aftercuring.

FIG. 8 is an enlarged, partially schematic isometric view of thelaminating station 620 configured in accordance with an embodiment ofthe invention. In one aspect of this embodiment, the laminating station620 includes a fiber placement machine 814 (shown schematically) movablysupported on a track beam 816. The track beam 816 can be part of a workplatform 822 positioned adjacent to the tool assembly 700 when the toolassembly 700 is parked in the laminating station 620. While notillustrated in detail in FIG. 8 for purposes of clarity, the fiberplacement machine 814 can include one or more payoff heads configured tocollimate multiple fiber tows 818. In addition, the fiber placementmachine 814 can further include supporting hardware (such as materialcreels, compaction rollers, etc.) typically used with multi-axis,gantry-mounted placement machines to dispense, clamp, cut, and restartfiber tows and/or other composite materials such as fabric, tapes,individual filaments, and other uni- and multidirectional preimpregnatedand non-preimpregnated composite materials and combinations thereof.

In operation, the fiber placement machine 814 moves back and forth alongthe track beam 816 laminating the collimated fiber tows 818 over theinnermost ply 721 as the tool assembly 700 rotates about thelongitudinal axis 707. The fiber placement machine 814 can include oneor more rollers or other suitable devices (not shown) for holding theinnermost ply 721 in place during application of the fiber tows 818 toavoid wrinkling of the innermost ply 721. The fiber placement machine814 can apply multiple plies in various patterns. For example, in oneembodiment, the fiber placement machine 814 can lay down plies on a−45/0/+45 degree bias to provide desired structural properties. In otherembodiments, other ply patterns and/or other orientations can be used toprovide other structural properties. In addition, hand lay-ups ofpreimpregnated fabric plies can also be applied over and in between towplies to provide additional strength around cut-outs and other localizedfeatures. In the foregoing manner, the fiber tows 818 together with theinnermost ply 721 form a continuous cylindrical skin or laminate 820extending around the plurality of stiffeners 730 (FIGS. 7A and 7B).

In the embodiment described above, the fiber placement machine 814applies fiber tows (e.g., carbon fiber tows preimpregnated with athermoset epoxy resin) to the laminate 820. Such fiber tows can havewidths from about 0.06 inch to about 0.50 inch (e.g., about 0.38 inch)after flattening by a compaction roller. In other embodiments, the fiberplacement machine can apply other types of tows, e.g., glass fiber tows,graphite fiber tows, and/or tows including other types of aramid fibersand resins.

In another embodiment, the fiber placement machine 814 can apply fibertape and/or slit fiber tape to the laminate 820 as the tool assembly 700rotates. The fiber tape can include a plurality of unidirectionalfibers, such as carbon fibers. The fibers can be interwoven with anothermaterial into a cloth tape, and/or the fibers can be held together by abacking paper that is removed prior to application.

In a further embodiment, the fiber placement machine 814 can applyindividual filaments to the laminate 820 in a filament winding process.In yet another embodiment, the fiber placement machine 814 can applyvarious combinations of the foregoing composite materials, as well ascomposite fabric sheets, to the laminate 820. The final layer ofmaterial applied to the laminate 820 can include a woven wire fabricthat provides both structural load carrying capability and lightningprotection. In the foregoing embodiments, the tool assembly 700 rotatesabout the longitudinal axis 707 as the fiber placement machine 814applies material. In other embodiments, however, the tool assembly 700can be rotationally fixed, and the fiber placement machine 814 can bemoved around the outside of the tool assembly 700 to apply material.After the final layer of material has been applied, the tool supportstructure 704 transports the tool assembly 700 from the laminatingstation 620 to the vacuum station 630 (FIG. 6) via the tracks 712.

FIG. 9 is an enlarged, partially schematic isometric view of the vacuumstation 630 configured in accordance with an embodiment of theinvention. In one aspect of this embodiment, the vacuum station 630includes opposing pressure pad supports 924 (identified individually asa first pressure pad support 924 a and a second pressure pad support 924b) movably positioned on opposite sides of the tracks 712. Each of thepressure pad supports 924 carries a corresponding pressure pad 926(identified individually as a first pressure pad 926 a and a secondpressure pad 926 b). In operation, the pressure pad supports 924 moveinwardly toward the tool assembly 700 to position the correspondingpressure pads 926 around the laminate 820 in a clam-shell configuration.In one embodiment, the pressure pads 926 can include conformablepressure pads and/or caul sheets configured to apply even pressure tothe laminate 820 during subsequent curing to produce a relatively smoothexterior surface. Once the pressure pads 926 have been installed on thelaminate 820, the pressure pad supports 924 are retracted and a vacuumbag (not shown) can be installed around the pressure pads 926 and thelaminate 820. After the vacuum bag has been evacuated, the tool assembly700 is lifted off of the tool support structure 704 and moved to thecuring station 640 (FIG. 6) via an overhead gantry beam 912. In otherembodiments, the vacuum bag can be omitted and the laminate 820 can becocured without prior evacuation.

FIG. 10 is an enlarged, partially schematic isometric view of the curingstation 640 configured in accordance with an embodiment of theinvention. In one aspect of this embodiment, the gantry beam 912 extendsfrom the vacuum station 630 into an autoclave 1050 positioned in thecuring station 640. The autoclave 1050 can include a door 1051 at eachend (identified individually as a first door 1051 a and a second door1051 b). The first door 1051 a retracts to allow the tool assembly 700to move into the autoclave 1050 on the gantry beam 912. Once the toolassembly 700 is positioned fully within the autoclave 1050, a gatesection 1013 of the gantry beam 912 moves out of the way to allow thefirst door 1051 a to move back into position. The temperature inside theautoclave 1050 is then elevated to cocure the laminate 820 and thestiffeners 730 (not shown). In one embodiment, the autoclave 1050 cancocure the laminate 820 and the stiffeners 730 using a standard 350° F.cure cycle. In other embodiments, other cure cycles can be useddepending on various factors such as material composition, thickness,etc. Once the parts have cooled, the second door 1051 b retracts asshown in FIG. 11, and the tool assembly 700 moves out of the autoclave1050 and on to the inspection station 650 via the gantry beam 912. Inother embodiments, the curing station 640 can include other systems formoving the tool assembly 700 in and out of the autoclave 1050. Suchsystems can include, for example, an autoclave cart, ground-based rails,etc.

FIG. 11 is an enlarged, partially schematic isometric view of theinspection station 650 configured in accordance with an embodiment ofthe invention. When the tool assembly 700 arrives at the inspectionstation 650, it is lowered from the gantry beam 912 onto a tool supportstructure 1104. Next, the laminate 820 is de-bagged and the pressurepads 926 (FIG. 9) are removed. The tool support structure 1104 can be atleast generally similar in structure and in function to the tool supportstructure 704 described above with reference to FIGS. 7A and 7B.Accordingly, the tool support structure 1104 can include a plurality ofrollers 1105 configured to rotate the tool assembly 700 about thelongitudinal axis 707.

In one aspect of this embodiment, the inspection station 650 includes aninspection machine 1160 movably supported adjacent to the tool supportstructure 1104. The inspection machine 1160 can be configured to moveback and forth along the length of the laminate 820 as the tool assembly700 rotates to inspect the structural integrity of the laminate 820. Inone embodiment, the inspection machine 1160 can include an ultrasonicinspection device for finding voids or disbonds in the laminate 820. Inother embodiments, other types of suitable inspection equipment known inthe art can be utilized to inspect the laminate 820. Such equipment mayinclude, for example, a pulse-echo inspection apparatus or athermographic inspection apparatus. Once the laminate 820 has been fullyinspected, the tool assembly 700 is again picked up by the gantry beam912 and moved to the trimming station 660 (FIG. 6).

FIG. 12 is an enlarged, partially schematic isometric view of thetrimming station 660 configured in accordance with an embodiment of theinvention. When the tool assembly 700 arrives at the trimming station660, it is lowered from the gantry beam 912 onto a tool supportstructure 1204. The tool support structure 1204 can be at leastgenerally similar in structure and in function to the tool supportstructures 704 and 1104 described above. Accordingly, the tool supportstructure 1204 can include a plurality of rollers 1205 configured torotate the tool assembly 700 about the longitudinal axis 707.

In one aspect of this embodiment, the trimming station 660 includes aCNC (computer numerically controlled) router 1270 and a CNC drillfixture 1272 movably supported adjacent to the tool support structure1204. Using determinate locator fixtures, the CNC router 1270 can beconfigured to form a plurality of window cutouts 1228 in the laminate820. The tool assembly 700 can rotate about the longitudinal axis 707 tofacilitate precise location of the window cutouts 1228. Similarly, theCNC drill fixture 1272 can be configured to drill a plurality offastener and/or assembly holes in the laminate 820 at this time. Afterthese trimming and drilling operations, barrel support rings (not shown)are positioned inside the laminate 820 to maintain the shell profilewhile the tool segments 706 (FIGS. 7A and 7B) are removed. The toolsegments 706 can then be returned to the stiffener loading station 610(FIG. 6) and prepared for the next fabrication cycle. After the toolsegments 706 have been removed, the tool support structure 1204transports the tool assembly 700 from the trimming station 660 to thefinal assembly station 670 (FIG. 6) via the floor tracks 712.

FIG. 13 is an enlarged, partially schematic isometric view of theassembly station 670 configured in accordance with an embodiment of theinvention. In one aspect of this embodiment, the final assembly station670 can include an internal work platform 1380 configured to support aninspection machine (not shown), such as a robotic ultrasonic inspectionmachine, for inspecting the structural integrity of the laminate 820from the interior surface. After this inspection, a plurality of framesections 1340 can be attached to the stiffeners 730 and/or the laminate820 from inside the tool assembly 700.

In one embodiment, the frame sections 1340 can be at least generallysimilar in structure and function to the frames 240 and/or the frames540 described above with reference to FIGS. 2A-B and 5A-B, respectively.In other embodiments, the frame sections 1340 can have other featuresor, alternatively, they can be omitted. The frame sections 1340 can belocated using the determinate assembly holes drilled previously at thetrimming station 660 (FIG. 12), and they can be attached using asemi-automated sealing and fastening process. The tool assembly 700 canrotate about the longitudinal axis 707 to facilitate installation of theframe sections 1340. In addition, a preassembled floor module (notshown) can be inserted, located, and attached to frame stub-outs at thistime. In a further aspect of this embodiment, the foregoingmanufacturing operations complete the basic structural assembly of thefuselage barrel section 110 to a point at which pre-assembled payloadsand interior kits can be installed. After that, the barrel section 110can be joined to adjacent barrel sections for final assembly of thefuselage 102 illustrated in FIG. 1.

FIGS. 14A-14C are cross-sectional end views illustrating various stagesof a method for bonding a stiffener 1430 to a laminate 1420 inaccordance with an embodiment of the invention. Referring first to FIG.14A, the uncured stiffener 1430 can be positioned in a tool 1406. Thestiffener 1430 can be a hat section stiffener (e.g., a hat sectionstiffener that is at least generally similar in structure and functionto the stiffeners 230 and 730 discussed above with reference to FIGS.2A-2B and FIGS. 7A-7B, respectively). In addition, the tool 1406 can beat least generally similar in structure and function to the tool segment706 described above with reference to FIGS. 7A-7B. After the stiffener1430 is positioned in the tool 1406, a tubular bladder 1480 supporting aportion of fabric 1482 (or tape, etc.) is positioned inside thestiffener 1430 so that the fabric 1482 contacts an interior surface 1432of the stiffener 1430 between opposing flange portions 1431 a and 1431b.

Referring next to FIG. 14B, once the bladder 1480 and the fabric 1482are positioned inside the stiffener 1430, composite materials arelaminated over the tooling segment 1406 to form a skin 1420 thatcontacts the flange portions 1431 and the fabric 1482. In one aspect ofthis embodiment, the skin 1420 can be at least generally similar instructure and function to the skin 220 and the laminate 820 describedabove with reference to FIGS. 2A-2B and FIG. 8, respectively.

Referring next to FIG. 14C, a compressible pad or caul sheet 1490 ispositioned over the skin 1420. Next, a vacuum bag 1492 is positionedaround the caul sheet 1490 and the tooling segment 1406. The spacebetween the vacuum bag 1492 and the bladder 1480 is then evacuated toapply an even pressure against the composite parts (i.e., the stiffener1430, the skin 1420, and the fabric 1482). The composite parts are thencocured at an elevated temperature while under vacuum. After curing, thestiffener/laminate combination is debagged and removed from the toolingsegment 1406.

In one embodiment of the method described above with reference to FIGS.14A-C, the stiffeners 1430 can be manufactured by laying-up one or moreplies of material directly into the tool 1406. In another embodiment,the stiffeners can be precured, or at least partially precured, beforeplacement in the tool 1406. When precured stiffeners are used, they canbe secondarily bonded to the skin 1420 with an adhesive during thesubsequent curing process.

One feature of the forgoing method is that the fabric 1482 serves as aninner doubler bonding the inner surface of the stiffener 1430 to anadjacent portion of the skin 1420 between the opposing flange portions1431. One advantage of this feature is that the fabric 1482 reduces thepeel stresses on the flange portions 1431. As a result, there is lesstendency for the stiffener 1430 to disbond from the skin 1420 under highhoop loads that may be encountered in service.

Various components described herein may be manufactured and/or assembledin accordance with the teachings of copending U.S. Provisional PatentApplication No. 60/559,911, entitled “STRUCTURAL PANELS FOR USE INAIRCRAFT FUSELAGES AND OTHER STRUCTURES,” and/or copending U.S. patentapplication Ser. No. 10/819,084, entitled “STRUCTURAL PANELS FOR USE INAIRCRAFT FUSELAGES AND OTHER STRUCTURES,” both of which were filed onApr. 6, 2004, and are incorporated herein in their entireties byreference.

Further, the subject matter of copending U.S. patent application Ser.No. 10/646,509, entitled “MULTIPLE HEAD AUTOMATED COMPOSITE LAMINATINGMACHINE FOR THE FABRICATION OF LARGE BARREL SECTION COMPONENTS,” filedAug. 22, 2003; Ser. No. 10/717,030, entitled “METHOD OF TRANSFERRINGLARGE UNCURED COMPOSITE LAMINATES,” filed Nov. 18, 2003; Ser. No.10/646,392, entitled “AUTOMATED COMPOSITE LAY-UP TO AN INTERNAL FUSELAGEMANDREL,” filed Aug. 22, 2003; Ser. No. 10/630,594, entitled “COMPOSITEFUSELAGE MACHINE,” filed Jul. 28, 2003; Ser. No. 10/646,316, entitled“UNIDIRECTIONAL, MULTI-HEAD FIBER PLACEMENT,” filed Aug. 22, 2003; Ser.No. 10/301,949, entitled “PARALLEL CONFIGURATION COMPOSITE MATERIALFABRICATOR,” filed Nov. 22, 2002; Ser. No. 10/799,306, entitled “SYSTEMSAND METHODS ENABLING AUTOMATED RETURN TO AND/OR REPAIR OF DEFECTS WITH AMATERIAL PLACEMENT MACHINE,” filed Mar. 12, 2004; Ser. No. 10/726,099,entitled “SYSTEMS AND METHODS FOR DETERMINING DEFECT CHARACTERISTICS OFA COMPOSITE STRUCTURE,” filed Dec. 2, 2003; Ser. No. 10/628,691,entitled “SYSTEMS AND METHODS FOR IDENTIFYING FOREIGN OBJECTS AND DEBRIS(FOD) AND DEFECTS DURING FABRICATION OF A COMPOSITE STRUCTURE,” filedJul. 28, 2003; and Ser. No. 10/822,538, entitled “SYSTEMS AND METHODSFOR USING LIGHT TO INDICATE DEFECT LOCATIONS ON A COMPOSITE STRUCTURE,filed Apr. 12, 2004, is incorporated herein in its entirety byreference. In addition, the subject matter of U.S. Pat. No. 6,168,358 isalso incorporated herein in its entirety by reference.

From the foregoing, it will be appreciated that specific embodiments ofthe invention have been described herein for purposes of illustration,but that various modifications may be made without deviating from thespirit and scope of the invention. For example, although the variousbarrel sections described above have been described in the context ofaircraft structures, in other embodiments, such sections can be used inother structural applications, such as space, water, and land vehicleapplications. Accordingly, the invention is not limited, except as bythe appended claims.

1. A composite aircraft panel manufacturing assembly comprising: a skin;a stiffener having first and second flange portions bonded to the skin,the stiffener further including an interior surface extending betweenthe first and second flange portions, wherein the interior surface isspaced apart from the skin and at least a portion of the interiorsurface faces toward the skin; a tubular bladder extendinglongitudinally between the interior surface of the stiffener and theskin; and at least one ply of fabric extending around the bladder andbonded to the interior surface of the stiffener and an adjacent portionof the skin between the first and second flange portions.
 2. Thecomposite aircraft panel manufacturing assembly of claim 1, wherein thefabric forms a tube contacting the interior surface of the stiffener andthe skin between the first and second flange portions.
 3. The compositepanel manufacturing assembly of claim 1, wherein the stiffener has ahat-shaped cross-section.
 4. The composite panel manufacturing assemblyof claim 1, wherein the skin includes a plurality of automaticallyplaced unidirectional fibers.